NPL 9615 AIRFOIL (npl9615-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: NPL 9615 AIRFOIL (npl9615-il) Reynolds number: 100,000 Max Cl/Cd: 43.29 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-npl9615-il-100000.txt Download as CSV file: xf-npl9615-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NPL 9615 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5484 0.10612 0.10076 -0.0116 1.0000 0.1121
-9.750 -0.5516 0.10210 0.09678 -0.0139 1.0000 0.1165
-9.500 -0.6082 0.09376 0.08855 -0.0278 1.0000 0.1192
-9.250 -0.5634 0.09264 0.08744 -0.0190 1.0000 0.1228
-9.000 -0.5605 0.08917 0.08400 -0.0196 1.0000 0.1270
-8.750 -0.5969 0.08257 0.07745 -0.0270 1.0000 0.1306
-8.500 -0.6642 0.07961 0.07418 -0.0281 1.0000 0.1325
-8.250 -0.6073 0.07482 0.06974 -0.0264 1.0000 0.1375
-8.000 -0.6180 0.07148 0.06634 -0.0263 1.0000 0.1433
-7.750 -0.6360 0.06741 0.06216 -0.0255 1.0000 0.1496
-7.500 -0.6219 0.06479 0.05961 -0.0240 1.0000 0.1557
-7.250 -0.6342 0.06150 0.05619 -0.0223 1.0000 0.1660
-7.000 -0.6255 0.05917 0.05386 -0.0204 1.0000 0.1747
-6.750 -0.6223 0.05628 0.05098 -0.0183 1.0000 0.1868
-6.500 -0.6209 0.05387 0.04855 -0.0159 1.0000 0.2032
-6.250 -0.6288 0.04084 0.03295 -0.0104 1.0000 0.0786
-6.000 -0.6176 0.03700 0.02855 -0.0071 1.0000 0.0716
-5.750 -0.6011 0.03444 0.02577 -0.0051 1.0000 0.0704
-5.500 -0.5848 0.03204 0.02303 -0.0029 1.0000 0.0701
-5.250 -0.5680 0.02991 0.02039 -0.0003 1.0000 0.0710
-5.000 -0.5475 0.02792 0.01843 0.0008 1.0000 0.0741
-4.750 -0.5286 0.02650 0.01660 0.0028 1.0000 0.0794
-4.500 -0.5094 0.02559 0.01559 0.0042 1.0000 0.0875
-4.250 -0.4874 0.02413 0.01426 0.0050 1.0000 0.0938
-4.000 -0.4651 0.02320 0.01307 0.0064 1.0000 0.0972
-3.750 -0.4423 0.02195 0.01197 0.0071 1.0000 0.1022
-3.500 -0.4201 0.02120 0.01116 0.0082 1.0000 0.1061
-3.250 -0.3982 0.02034 0.01035 0.0092 1.0000 0.1098
-3.000 -0.3775 0.01968 0.00981 0.0103 1.0000 0.1151
-2.750 -0.3574 0.01916 0.00929 0.0115 1.0000 0.1205
-2.500 -0.3389 0.01859 0.00887 0.0128 1.0000 0.1280
-2.250 -0.3205 0.01813 0.00848 0.0140 1.0000 0.1369
-2.000 -0.3017 0.01779 0.00816 0.0152 1.0000 0.1477
-1.750 -0.2828 0.01733 0.00783 0.0162 1.0000 0.1661
-1.500 -0.2694 0.01467 0.00785 0.0187 1.0000 0.6742
-1.250 -0.2111 0.01538 0.00924 0.0180 1.0000 0.9470
-1.000 -0.0048 0.01642 0.00976 -0.0122 1.0000 1.0000
-0.750 -0.0335 0.01646 0.00982 -0.0053 1.0000 1.0000
-0.500 0.0168 0.01654 0.00978 -0.0112 0.9899 1.0000
-0.250 0.0763 0.01649 0.00965 -0.0184 0.9763 1.0000
0.000 0.1338 0.01639 0.00948 -0.0250 0.9627 1.0000
0.250 0.1949 0.01618 0.00924 -0.0320 0.9499 1.0000
0.500 0.2575 0.01581 0.00886 -0.0388 0.9363 1.0000
0.750 0.3081 0.01544 0.00849 -0.0431 0.9171 1.0000
1.000 0.3491 0.01512 0.00817 -0.0455 0.8959 1.0000
1.250 0.3835 0.01484 0.00786 -0.0462 0.8737 1.0000
1.500 0.4079 0.01467 0.00766 -0.0450 0.8463 1.0000
1.750 0.4309 0.01445 0.00738 -0.0432 0.8190 1.0000
2.000 0.4529 0.01430 0.00714 -0.0414 0.7918 1.0000
2.250 0.4763 0.01424 0.00696 -0.0397 0.7669 1.0000
2.500 0.4975 0.01432 0.00697 -0.0380 0.7391 1.0000
2.750 0.5190 0.01440 0.00697 -0.0363 0.7128 1.0000
3.000 0.5401 0.01449 0.00699 -0.0346 0.6863 1.0000
3.250 0.5621 0.01460 0.00698 -0.0330 0.6618 1.0000
3.500 0.5828 0.01475 0.00711 -0.0314 0.6351 1.0000
3.750 0.6041 0.01490 0.00718 -0.0298 0.6096 1.0000
4.000 0.6247 0.01509 0.00733 -0.0281 0.5822 1.0000
4.250 0.6452 0.01531 0.00744 -0.0264 0.5539 1.0000
4.500 0.6647 0.01556 0.00763 -0.0245 0.5218 1.0000
4.750 0.6838 0.01585 0.00782 -0.0226 0.4874 1.0000
5.000 0.7021 0.01622 0.00807 -0.0205 0.4494 1.0000
5.250 0.7195 0.01669 0.00834 -0.0184 0.4093 1.0000
5.500 0.7366 0.01726 0.00869 -0.0163 0.3699 1.0000
5.750 0.7540 0.01793 0.00913 -0.0143 0.3358 1.0000
6.000 0.7722 0.01860 0.00970 -0.0125 0.3068 1.0000
6.250 0.7910 0.01939 0.01028 -0.0109 0.2846 1.0000
6.500 0.8100 0.02009 0.01096 -0.0094 0.2640 1.0000
6.750 0.8288 0.02085 0.01164 -0.0078 0.2451 1.0000
7.000 0.8469 0.02159 0.01234 -0.0062 0.2265 1.0000
7.250 0.8642 0.02227 0.01294 -0.0045 0.2086 1.0000
7.500 0.8810 0.02301 0.01362 -0.0028 0.1910 1.0000
7.750 0.8968 0.02374 0.01439 -0.0009 0.1728 1.0000
8.000 0.9127 0.02465 0.01517 0.0009 0.1554 1.0000
8.250 0.9271 0.02563 0.01615 0.0030 0.1368 1.0000
8.500 0.9414 0.02677 0.01723 0.0051 0.1196 1.0000
8.750 0.9567 0.02802 0.01844 0.0070 0.1058 1.0000
9.000 0.9743 0.02950 0.01994 0.0085 0.0959 1.0000
9.250 0.9917 0.03090 0.02140 0.0100 0.0882 1.0000
9.500 1.0108 0.03221 0.02265 0.0112 0.0825 1.0000
9.750 1.0281 0.03389 0.02450 0.0126 0.0777 1.0000
10.000 1.0451 0.03551 0.02624 0.0140 0.0736 1.0000
10.250 1.0611 0.03736 0.02825 0.0154 0.0702 1.0000
10.500 1.0813 0.03908 0.02992 0.0160 0.0672 1.0000
10.750 1.0865 0.04158 0.03294 0.0187 0.0651 1.0000
11.000 1.1019 0.04359 0.03502 0.0199 0.0630 1.0000
11.250 1.1094 0.04639 0.03812 0.0218 0.0620 1.0000
11.500 1.1049 0.04965 0.04185 0.0249 0.0615 1.0000
11.750 1.0960 0.05303 0.04565 0.0279 0.0613 1.0000
12.000 1.0812 0.05641 0.04936 0.0311 0.0612 1.0000
12.250 1.0616 0.05992 0.05316 0.0340 0.0613 1.0000
12.500 1.0398 0.06390 0.05740 0.0357 0.0616 1.0000
12.750 1.0167 0.06844 0.06216 0.0359 0.0619 1.0000
13.000 0.9938 0.07358 0.06749 0.0348 0.0622 1.0000
13.250 0.9728 0.07920 0.07326 0.0328 0.0625 1.0000
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Polar data table (+)
Polar graphs
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